This invention relates generally to gas turbine engines, and in particular, to a cooled flow path surface region.
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000xc2x0 F., considerably higher than the melting temperatures of the metal parts of the engine, which are in contact with these gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling air to the outer surfaces of the metal parts through various methods. The metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are metal parts forming combustors and parts located aft of the combustor including turbine blades, turbine vanes and exhaust nozzles.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at metal surface temperatures of up to 2100xc2x0-2200xc2x0 F.
The metal temperatures can be maintained below melting levels with current cooling techniques by using a combination of improved cooling designs and insulating thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades essentially have intricate serpentine passageways within structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from the hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a TBC is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts within engines to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component. TBCs have also been used in combination with film cooling techniques wherein an array of fine holes extends from the hollow core through the TBC to provide cooling air onto the outer surface of the TBC.
TBCs are well-known ceramic coatings, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used in the substrates. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the thermal barrier coating. The bond coat may be made of a nickel-containing overlay alloy, such as a NiCrAlY or a NiCoCrAIY, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide, whose surface oxidizes to a protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and the overlying TBC are frequently referred to as a thermal barrier coating system.
Multi layer coatings are known in the art. For example, U.S. Pat. No. 5,846,605 to Rickerby, et al., is directed to a coating having a plurality of alternate layers having different structures that produce a plurality of interfaces. The interfaces provide paths of increased resistance to heat transfer to reduce thermal conductivity. A bond coat overlying a metallic substrate is bonded to a TBC. The TBC comprises a plurality of layers, each layer having columnar grains, the columnar grains in each layer extending substantially perpendicular to the interface between the bond coating and metallic substrate. The structure is columnar to ensure that the strain tolerance of the ceramic TBC is not impaired. The difference in structure of the layers is the result of variations in the microstructure and/or density/coarseness of the columnar grains of the ceramic.
U.S. Pat. No. 5,705,231 to Nissley et al. is directed to a segmented abradable ceramic coating system having enhanced abradability and erosion resistance. A segmented abradable ceramic coating is applied to a bond coat comprising three ceramic layers that are individually applied. There is a base coat foundation layer, a graded interlayer, and an abradable top layer. The coating is characterized by a plurality of vertical microcracks.
U.S. Pat. No. 4,503,130 to Bosshart et al. is directed to coatings having a low stress to strength ratio across the depth of the coating. Graded layers of metal/ceramic material having increasing ceramic composition are sequentially applied to the metal substrate under conditions of varied substrate temperature.
U.S. Pat. No. 6,045,928 to Tsantrizos et al. is directed to a TBC comprising an MCrAlY bond coat and a dual constituent ceramic topcoat. The topcoat comprises a monolithic zirconia layer adjacent to the bond coat, a monolithic layer of calciasilica representing the outer surface of the TBC and a graded interface between the two to achieve good adhesion between the two constituents to achieve an increased thickness of the topcoat, thereby, providing for a greater temperature drop across the TBC system.
U.S. Pat. No. 4,576,874 to Spengler et al. is directed to a coating to increase resistance to spalling and corrosion. The coating is not intended to be a thermal barrier coating. A porous ceramic is applied over a MCrAlY bond coat and a dense ceramic is then applied over the porous ceramic. The porous portion is a transition zone to allow for differences in thermal expansion and provides little thermal insulation.
Improved environmental resistance to destructive oxidation and hot corrosion is desirable. In some instances, the alloying elements of the bond coat can interdiffuse with the substrate alloy and consume walls of the turbine airfoils, i.e., reduce load carrying capability. This interdiffusion also reduces environmental resistance of the coating. Even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling; because reducing the demand for cooling is also well known to improve overall engine operating efficiency.
While superalloys coated with thermal barrier coating systems do provide substantially improved performance over uncoated materials, there remains room for improvement. Film cooling is achieved by passing cooling air through discrete film cooling holes, typically ranging from 0.015xe2x80x3 to about 0.030xe2x80x3 in hole diameters. The film cooling holes are typically drilled with laser or EDM or ES machining. Due to mechanical limitations, each film hole has an angle ranging from 20xc2x0 to 90xc2x0 relative to the external surface. Therefore, each film jet exits from the hole with a velocity component perpendicular to the surface. Because of this vertical velocity component and a flow circulation around each jet due to the jet mixing, each film jet will have a tendency to lift or blow off from the external surface and mix with the hot exhaust gases, resulting in poor film cooling effectiveness.
Thus, there is an ongoing need for an improved thermal barrier coating system, wherein the environmental resistance and long-term stability of the thermal barrier coating system is improved so that higher engine efficiencies can be achieved. The bond coat temperature limit is critical to the TBC""s life and has had an upper limit of about 2100xc2x0 F. Once the bond coat exceeds this temperature, the coating system will quickly deteriorate, due to high temperature mechanical deformation and oxidation, as well as interdiffusion of elements with the substrate alloy. The coating system can separate from the substrate exposing the underlying superalloy component to damage from the hot gasses.
What is needed are improved designs that will allow turbine engine components to run at higher operating temperatures, thus improving engine performance without additional cooling air. It is desirable to have a system that can take advantage of the thermal insulation provided by TBC. The present invention fulfills this need, and further provides related advantages.
The present invention provides a method for cooling the flow path surface region of an engine component used in a gas turbine engine comprising the steps of channeling apertures in a substrate to a diameter of about 0.0005xe2x80x3 to about 0.02xe2x80x3 to allow passage of cooling fluid from a cooling fluid source; applying a bond coat of about 0.0005xe2x80x3 to about 0.005xe2x80x3 in thickness to the substrate such that the bond coat partially fills the channels; applying a porous inner TBC layer of at least about 0.003xe2x80x3 in thickness to the bond coat, such that the TBC fills the channels; applying an intermediate ceramic layer that is more dense than the inner TBC layer on top of the porous TBC; applying an outer TBC layer that is more dense than the inner TBC layer over the intermediate layer; and, passing cooling fluid from a cooling fluid source through the channel into the porous TBC. Because the channel exit is filled with porous TBC material, cooling fluid flows through the porous passageways into the inner TBC layer. Although the passageways are interconnected and provide a plurality of tortuous routes, the increased density of the TBC in the intermediate layer provides a resistance to flow of the cooling fluid and effectively causes the cooling fluid to more efficiently spread through the TBC in the inner layer before exiting at the outer surface.
The present invention further comprises both the cooled flow path surface regions formed by the foregoing methods and the turbine component with the ceramic layers for cooling the component.
In a different embodiment, the present invention comprises a cooling channel having a first and second end, the first end terminating in an exit orifice located on the surface of a substrate, the second end connecting to a cooling circuit contained within a turbine engine component. The cooling channel preferably has a substantially circular diameter of about 0.002xe2x80x3 to about 0.008xe2x80x3. Applied to the substrate is a bond coat of about 0.0005xe2x80x3 to about 0.005xe2x80x3 in thickness that partially fills the exit orifice and first channel end. Applied to the bond coat is a porous inner TBC layer of at least about 0.003xe2x80x3 in thickness such that the porous TBC fills the remainder of the exit orifice. An intermediate ceramic layer that is more dense than the porous inner TBC is applied on top of the porous inner TBC, and an outer TBC layer that is less porous than the inner TBC layer is then applied on top of the intermediate layer.
An advantage of the present invention is that the multi-layered TBC system forms cooling paths adjacent to the substrate surface to provide efficient cooling for both the substrate and the bond coat by allowing heat to be removed from the article.
Still another advantage of the present invention is that the outer flow path surface region of the coated gas turbine component is actively cooled through transpiration cooling inside the TBC. Transpiration cooling inside the TBC will lower the TBC temperature, and allow a greater thermal gradient between the hot exhaust gas stream and the bond coat. By removing heat from this region by transpiration, the integrity of the bond coat can be maintained at higher engine firing temperatures, resulting in a more efficient usage of cooling fluid to achieve a higher turbine engine efficiency and performance.
Another advantage of the present invention is that because the TBC is processed to have a varying density, and hence variable porosity, cooling fluid is able to flow through the inner TBC passageways and further spread through the outer TBC layer providing transpiration cooling before exiting to the outer TBC surface or flow through discrete holes providing film cooling, or both.
Still another advantage of the present invention is that the cooling channel exit orifices or holes, being filled, have more flow resistance than open holes and, therefore, provide a more effective cross-sectional hole area compared to unfilled larger holes for flowing the same amount of cooling fluid, resulting in more efficient heat transfer.
Yet another advantage is that the characteristics of many smaller passageways through the TBC will result in a much larger heat transfer area for cooling the bond coat and the substrate than that provided by fewer larger holes having the same total flow cross-section area.
Still another advantage is that improved cooling reduces or eliminates sintering of the ceramic top coat so that low thermal conductivity is maintained in the TBC.